High-energy to low-energy orbit transfer vehicle

ABSTRACT

The excess space and weight capacity of a conventional launch vehicle for a high-energy orbit, such as GEO, is used to deploy satellites to a low-energy orbit, such as LEO. In a preferred embodiment, an orbit-transfer vehicle provides the navigation, propulsion, and control systems required to transport a payload satellite from a high-energy-transfer orbit, such as GTO, to a predetermined low-energy orbit. Upon entering the low-energy orbit, the payload satellite is released from the orbit-transfer vehicle. To reduce the fuel requirements for this deployment via the orbit-transfer vehicle, a preferred embodiment includes aerobraking to bring the satellite into a low-earth orbit. In a preferred embodiment of this method of deployment, the provider of the orbit-transfer vehicle identifies and secures available excess capacity on launch vehicles, and allocates the excess capacity to the satellites requiring low-earth orbit deployment, thereby providing a deployment means that is virtually transparent to the purchaser of this deployment service.

CROSS-REFERENCE TO RELATED APPLICATIONS

This is a Division of U.S. patent application Ser. No. 09/350,813, filedJul. 9, 1999, now U.S. Pat. No. 6,286,787.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to the field of aerospace, and generally to thedeployment of small satellites in low-earth orbit (LEO).

2. Description of Related Art

Satellites, because of their unobstructed fields of view of large areasof the earth, are often the preferred technical solutions to a varietyof communications and monitoring problems. The high cost of satellitedeployment, however, often precludes their use from a cost-efficiencyviewpoint. In order to distribute the high costs of deployment among alarge number of applications or users, the majority of deployedsatellites are those that handle a multitude of tasks, or a multitude ofcustomers for the same task. To minimize the loss of available accesstime to or from the satellite from or to the location on earth beingserviced by a satellite, many satellites are placed in geosynchronousorbit. A geosynchronous orbit tracks the revolution of the earth, sothat the satellite appears to be fixed over the same area of the earth,thus providing continual access to that area. Due to the physicsinvolved, a geosynchronous orbit is approximately 36,000 kilometersabove the earth. To provide reliable communications over this longdistance, a geosynchronous satellite requires highly sensitive receiversand/or highly powerful transmitters. Because of the aforementionedeconomic and technical requirements, geosynchronous satellites generallyweigh a ton or more, and cost (in 1999) hundreds of millions of dollarsto deploy to the selected geosynchronous orbit.

As contrast to large geosynchronous satellites, the use of smallsatellites at low-earth orbit (LEO) are becoming increasingly common.Copending U.S. patent applications “SATELLITE COMMUNICATION SYSTEM WITHA SWEEPING HIGH-GAIN ANTENNA”, U.S. Ser. No. 09/045,971, filed Mar. 21,1998 for Fleeter et al; “LOW-COST SATELLITE COMMUNICATION SYSTEM”, U.S.Ser. No. 09/045,970, filed Mar. 21, 1998 for Fleeter et al; “IN SITUREMOTE SENSING”, U.S. Ser. No. 09/130,854, filed Aug. 7, 1998 forRichard Fleeter; “RF INSPECTION SATELLITE”, U.S. Ser. No. 09/267,942,filed Mar. 11, 1999 for Hanson et al, illustrate the use of low costsatellites for a variety of applications, and are incorporated byreference herein. Low-earth orbits are typically hundreds of miles abovethe earth, rather than thousands of miles. Because of their order ofmagnitude closer proximity to earth, satellites in low-earth orbitrequire significantly less communicating and monitoring power andsensitivity than the satellites in geosynchronous orbit. Because theyare not stationary above any location on the earth, multiple satellitesin low-earth orbit are required to provide continuous coverage of aparticular area on earth. Because multiple satellites are required inlow-earth orbit to provide continuous coverage, a low-earth orbitsatellite system is particularly well suited to applications that employlow cost satellites. As advances continue to be made in electroniccircuit density and efficiency, the number of communication andmonitoring applications that can be embodied in small, low costsatellites continues to increase.

Deployment of a small, less than five hundred pound, satellite intolow-earth orbit typically costs, in 1999 dollars, between seven and tenmillion dollars. Because a plurality of satellites is required toprovide continuous coverage of an area, the overall cost of deployingconstellations of low-earth orbit satellites can often amount tohundreds of millions of dollars.

BRIEF SUMMARY OF THE INVENTION

It is an object of this invention to provide a lower cost means fordeploying a satellite into low-earth orbit. It is a further object ofthis invention to provide a method for economically brokering thedeployment of a satellite into low-earth orbit. It is a further objectof this invention to provide an orbit-transfer vehicle to effectivelydeploy small satellite systems to low-earth orbit.

A launch of geosynchronous satellites typically includes one or twolarge, multi-ton, satellites that are deployed at the geosynchronousaltitude of 36,000 kilometers via a large multistage rocket system, suchas an Ariane system. Typically, after allocating the available space andweight capabilities of the rocket system to the primary payload of theone or two large satellites, some excess space and weight allocationremains. For example, if an Ariane launch vehicle can accommodate fourtons, and the primary payload satellites are 1½ and 2 tons each, thelaunch vehicle has an excess capacity of a half ton. Because themarginal cost of adding one or two small satellites is minimal, thisexcess space or weight capacity can be brokered for the deployment ofsmall satellites at substantially less cost than the primary payload,often less than a quarter of the cost per pound charged to the primarypayload satellites.

The expressed objects of this invention, and others, are achieved byproviding a means of utilizing the excess space and weight capacity thatis typical of a launch of large geosynchronous satellites to deploysmall satellites at a low-earth orbit. Specifically, this inventionprovides a method of deployment of small satellite systems to low-earthorbit from a geosynchronous-transfer launch vehicle. In a preferredembodiment, an orbit-transfer vehicle provides the navigation,propulsion, and control systems required to transport a payloadsatellite from a geosynchronous-transfer orbit (GTO) to a predeterminedlow-earth orbit (LEO). Upon entering low-earth orbit, the payloadsatellite is deployed from the orbit-transfer vehicle. To reduce thefuel requirements for this deployment via the orbit-transfer vehicle, apreferred embodiment includes aerobraking to bring the satellite into alow-earth orbit. In a preferred embodiment of this method of deployment,the provider of the orbit-transfer vehicle identifies and securesavailable excess capacity on geosynchronous-transfer launch vehicles,and allocates the excess capacity to the satellites requiring low-earthorbit deployment, thereby providing a deployment means that is virtuallytransparent to the purchaser of this deployment service.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in further detail, and by way of example,with reference to the accompanying drawings wherein:

FIG. 1 illustrates a conventional deployment of a geosynchronoussatellite via a geosynchronous-transfer launch vehicle.

FIG. 2 illustrates an example deployment of a low-earth orbit (LEO)satellite using a geosynchronous-transfer launch vehicle in accordancewith this invention.

FIG. 3 illustrates an example method of facilitating the use of excesscapacity of a geosynchronous-transfer launch vehicle for deploying alow-earth orbit satellite in accordance with this invention.

FIG. 4 illustrates an example orbit-transfer vehicle in accordance withthis invention.

FIG. 5 illustrates an alternative example orbit-transfer vehicle inaccordance with this invention.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 illustrates a conventional deployment of a geosynchronoussatellite via a geosynchronous-transfer launch vehicle. Note that thefigures are presented herein for illustration purposes. Although therelative size of the illustrated orbits are approximately to scalerelative to the illustrated size of the earth 130, the specific detailsof the orbits should not be interpreted as being representative. Ageosynchronous-transfer launch vehicle (not shown) containing one ormore satellites for deployment in geosynchronous orbit 150 is launched101 from the earth 130. The geosynchronous-transfer launch vehicle isconfigured to achieve a geosynchronous-transfer orbit 120 whose apogee121 is tangent to the geosynchronous orbit 150. Some time afterachieving the geosynchronous-transfer orbit 120, the one or moregeosynchronous satellites (not shown) are released from thegeosynchronous-transfer launch vehicle. All items that are released bythe geosynchronous-transfer launch vehicle have the same velocity as thegeosynchronous-transfer launch vehicle when they are released, andtherefore continue to travel in the geosynchronous-transfer orbit 120.Each geosynchronous satellite contains a means for applying thrust 102,preferably at or near apogee 121, to enter an intended tangentialgeosynchronous orbit 150.

FIG. 2 illustrates an example method of deploying a low-earth orbit(LEO) satellite using a geosynchronous-transfer launch vehicle inaccordance with this invention. As noted above, a geosynchronous orbit150 is located 36,000 kilometers above the earth 130. A low-earth orbit110, on the other hand, is located well within 1500 kilometers of theearth 130. The Space Shuttle, for example, typically orbits the earth atan altitude between 160 and 600 kilometers. For ease of reference, theterm geosynchronous-transfer launch vehicle is used to define a launchvehicle that deploys its payload at an elevation above the earth that issubstantially higher than a low-earth orbit, because the most common useof a high-altitude launch vehicle is for geosynchronous satellitedeployment.

In accordance with this invention, an orbit-transfer vehicle (not shown)containing the low-earth orbit satellite (not shown) is launched 101from the earth 130 via a conventional geosynchronous-transfer launchvehicle (not shown), such as an Ariane. The geosynchronous-transferlaunch vehicle releases the orbit-transfer vehicle intogeosynchronous-transfer orbit 120. Some time after release, preferablyat the apogee 121 of the geosynchronous-transfer orbit 120, theorbit-transfer vehicle fires 103 its integral propulsion device (notshown) to force itself, and the attached low-earth orbit satellite, outof the geosynchronous-transfer orbit 120, and begins a forced descenttoward the earth 130. Another firing 104 of the propulsion device canthereafter be used to place the orbit-transfer vehicle directly into alow-earth orbit 110, but such a direct geosynchronous-transfer orbit tolow-earth orbit transfer will require a substantial amount of fuel toreduce the kinetic energy of the orbit-transfer vehicle sufficiently toremain in the low-earth orbit 110.

In accordance with an aspect of this invention, aerobraking is used tofacilitate the geosynchronous-transfer orbit to low-earth orbittransfer. Aerobraking uses the friction of the earth's atmosphere toreduce the kinetic energy of a spacecraft. Illustrated in FIG. 2, thefiring 103 of the propulsion device provides a thrust to force theorbit-transfer vehicle close to the earth 130, at 113. In a preferredembodiment, the firing 103 of the propulsion device is controlled toprovide a perigee 113 to within a hundred kilometers above the earth130. At this nominal hundred kilometer altitude, the atmosphere of theearth is sufficiently dense so as to impart a frictional force thatreduces the kinetic energy of the orbit-transfer vehicle, and itsattached low-earth orbit satellite. This reduction in kinetic energyresults in an apogee 114 of the orbit-transfer vehicle that is less thanits original geosynchronous-transfer orbit apogee 121. Upon achievingapogee 114, the orbit-transfer vehicle is drawn toward the earth 130again, and reaches perigee 115, also within 100 kilometers of the earth130. The friction of the earth's atmosphere at this low altitude againreduces the kinetic energy of the orbit-transfer vehicle, and theresultant apogee 116 is less than the prior apogee 114. Subsequenttraversals 117 of the earth's atmosphere will continue to reduce thekinetic energy of the orbit-transfer vehicle, further lowering eachsubsequent apogee 118.

When sufficient kinetic energy is removed from the orbit-transfervehicle, the decreasing apogee 111 of the orbit-transfer vehiclesubstantially approaches the altitude of the intended low-earth orbit110. At this apogee 111, the orbit-transfer vehicle effects the firing105 of the integral propulsion device to force an ascent of theorbit-transfer vehicle so as to raise the perigee of the orbit-transfervehicle, and attached low-earth orbit satellite, beyond the earth'satmosphere, thereby preventing further decreases of apogee. Thepreferred firing 105 places the orbit transfer vehicle, and attachedlow-earth orbit satellite, into a symmetric low-earth orbit 110, with aperigee that is substantially equal to the apogee 111. In a preferredembodiment, over one hundred aerobraking orbits are made before firingthe integral propulsion device at 105. Typically, the orbit-transfervehicle releases the low-earth orbit satellite at this low-earth orbit110, and thereafter the low-earth orbit satellite operatesindependently, as it would have, had it been launched directly from theearth 130 to the low-earth orbit 110. That is, although the low-earthorbit satellite may travel hundreds of thousands of miles to reach anorbit 110 that is only a few hundred miles above the earth's surface,its operation is substantially independent of this rather circuitousdeployment scheme.

Because the deployment methods in accordance with this invention can beeffected without affecting the satellite payload, this inventionprovides a means for brokering lower cost services for the deployment oflow-earth orbit satellites. Generally, a satellite is used as acomponent of a ground-based system, such as a communications network, aresearch facility, and the like. The user, or owner, of the satellite isnot necessarily fluent in the intricacies of rocketry and orbitalmechanics. A satellite deployment broker provides the interface servicesand support between the owner of the satellite and the provider of alaunch vehicle. In accordance with the principles of this invention, asatellite deployment broker can extend the range of potential low-earthorbit satellite launch vehicle providers to includegeosynchronous-transfer launch vehicle providers having excess capacity.FIG. 3 illustrates, for example, a flow diagram for allocating launchservices for low-earth orbit satellite deployments. At 210, thesatellite(s) requirements are determined, including the required orbitparameters, the size and weight of the satellite, and so on. At 220, thecost of a conventional low-earth orbit launch that satisfies therequirements are determined or estimated. At 230, the availability ofexcess capacity on scheduled geosynchronous orbit launch vehicles isdetermined, and a cost is negotiated for using this excess capacity. Theoverall cost of deploying the low-earth orbit satellite via ageosynchronous-transfer launch is the cost of using the excess launchcapacity of the geosynchronous-transfer launch vehicle plus the cost ofthe orbit-transfer vehicle for transporting the satellite from thegeosynchronous orbit to the low-earth orbit, as determined at 240. If,at 250, the overall cost of the geosynchronous-transfer launch andorbit-transfer is less than the conventional low-earth orbit launch, thedeployment is effected by attaching 260 the satellite to theorbit-transfer vehicle and launching 270 the orbit-transfer vehicle withsatellite via the geosynchronous-transfer launch vehicle. Thereafter,the orbit-transfer vehicle effects the deployment 280 of the satelliteto a low-earth orbit as discussed above. If, at 250, the cost of theconventional low-earth orbit launch is less expensive than thegeosynchronous-transfer launch, the satellite is deployed 290 via theconventional low-earth orbit launch. Note that a deployment of aconstellation of satellites to low-earth orbits in accordance with thisinvention can involve a combination of low-earth orbit andgeosynchronous-transfer launches, depending primarily on theavailability and cost of excess capacity on scheduledgeosynchronous-transfer launch vehicles.

FIG. 4 illustrates an example orbit-transfer vehicle 300 in accordancewith this invention. The example orbit-transfer vehicle 300 includes anadapter element 310 that provides a conventional means 315 for securingthe vehicle 300 to the geosynchronous-transfer launch vehicle (notshown), a body element 320 that provides a cavity for holding a fuelcell 340, and an attitude determination and control system 330 thatprovides the navigation, propulsion, and control systems required totransport the satellite from a geosynchronous-transfer orbit to apredetermined low-earth orbit. The adapter element 310 and body 320 in apreferred embodiment include shielding to enable the orbit-transfervehicle 300 to withstand the heat that is induced by the atmosphericfriction during aerobraking. The elevation at perigee 113, 115, 117determines the required degree of shielding. Correspondingly, theelevation at perigee 113, 115, 117 determines the number of aerobrakingorbits required to provide a sufficient reduction in the kinetic energyof the orbit-transfer vehicle 300 to achieve a low-earth orbit, giventhe capacity of the fuel cell 340. Jets 325 on the body element 320effect the thrust required to effect the orbit-transfer, under thecontrol of the attitude determination and control system 330. Asatellite 380 is designed to be mounted within the body 320, and isreleased from the orbit-transfer vehicle 300 when the appropriatelow-earth orbit is achieved. Alternative arrangements will be evident toone of ordinary skill in the art in view of this disclosure. Forexample, FIG. 5 illustrates an orbit-transfer vehicle 400 for use with asatellite 480 having an integral body 420 and conventional means 315 forsecuring the satellite 300 to a launch vehicle (not shown).

The foregoing merely illustrates the principles of the invention. Itwill thus be appreciated that those skilled in the art will be able todevise various arrangements which, although not explicitly described orshown herein, embody the principles of the invention and are thus withinits spirit and scope. For example, the orbit-transfer components 430 canbe built into the satellite 480, thereby eliminating the need to detachthe components 430 upon achievement of the low-earth orbit.

I claim:
 1. A method of deploying a payload satellite into a targetorbit having an associated target-orbit energy-level, comprising:placing the payload satellite on a launch vehicle, launching the launchvehicle, deploying the payload satellite from the launch vehicle with anassociated kinetic energy, the kinetic energy associated with thepayload satellite being substantially greater than the target-orbitenergy-level associated with the target orbit, and maneuvering thepayload satellite into the target orbit via a reduction in the kineticenergy associated with the payload satellite, to an energy levelcorresponding to the target-orbit energy-level.
 2. The method of claim1, wherein the reduction in kinetic energy includes: aerobraking thepayload satellite.
 3. The method of claim 2, further including attachingthe payload satellite to an orbit-transfer vehicle before the payloadsatellite is placed on the launch vehicle, and wherein theorbit-transfer vehicle effects the aerobraking of the payload satellite.4. The method of claim 2, wherein the aerobraking includes applyingthrust to the payload satellite so as to force the payload satellite toenter a portion of atmosphere above the earth when the payload satellitesubstantially reaches perigee.
 5. The method of claim 4, wherein theaerobraking further includes applying thrust to the payload satellite soas to force the payload satellite to an orbit beyond the atmosphereabove the earth when the payload satellite substantially reaches apogeeat an elevation corresponding to the low-earth orbit.
 6. The method ofclaim 1, further including attaching the payload satellite to anorbit-transfer vehicle before the payload satellite is placed on thelaunch vehicle, and wherein the step of maneuvering the payloadsatellite is effected by at least one of: the payload satellite, and theorbit-transfer vehicle.
 7. The method of claim 6, further includingdetaching the payload satellite from the orbit-transfer vehicle.
 8. Amethod of facilitating the deployment of a payload satellite into atarget orbit having an associated target-orbit energy-level, comprising:identifying an excess capacity on a scheduled launch vehicle having anassociated transfer orbit that has an associated orbit-transfer energylevel that is substantially higher than the target-orbit energy-level,facilitating an attachment of the payload satellite to the launchvehicle, facilitating a deployment of the payload satellite into thetransfer orbit via the launch vehicle, the payload satellite therebyhaving a kinetic energy corresponding to the orbit-transfer energylevel, and facilitating a maneuvering of the payload satellite to thetarget orbit via a substantial reduction in the kinetic energy of thepayload satellite, from the orbit-transfer energy level to thetarget-orbit energy-level.
 9. The method of claim 8, further includingfacilitating an aerobraking of the payload satellite so as to achievethe target orbit.
 10. The method of claim 8, wherein the attachment ofthe payload satellite to the launch vehicle is via: an attachment of thepayload satellite to an orbit-transfer vehicle, and an attachment of theorbit-transfer vehicle to the launch vehicle.
 11. The method of claim 8,wherein the maneuvering of the payload satellite to the target orbit isvia an orbit-transfer vehicle.
 12. An orbit-transfer vehicle thatfacilitates the deployment of a payload satellite into a target orbit,comprising an attitude determination and control system that isconfigured to maneuver the orbit-transfer vehicle and the payloadsatellite from a launch orbit to the target orbit the launch orbithaving an associated launch-orbit-energy that is substantially greaterthan a target-orbit-energy associated with the target orbit, wherein theorbit-transfer vehicle is configured to facilitate: an attachment of thepayload satellite to the orbit-transfer vehicle, and an attachment ofthe orbit-transfer vehicle to a launch vehicle that is configured todeploy the orbit-transfer vehicle with attached payload satellite to thetarget orbit.
 13. The orbit-transfer vehicle of claim 12, furtherincluding the payload satellite.
 14. The orbit-transfer vehicle of claim12, wherein the attitude determination and control system is configuredto provide a thrust that is sufficient to force the orbit-transfervehicle from the launch orbit into a portion of atmosphere to effect adecrease in kinetic energy of the orbit-transfer vehicle.
 15. Theorbit-transfer vehicle of claim 14, wherein the attitude determinationand control system is further configured to apply thrust to the payloadsatellite so as to force the payload satellite beyond the atmospherewhen the payload satellite substantially reaches apogee at an elevationcorresponding to the target orbit.
 16. The orbit-transfer vehicle ofclaim 12, wherein the orbit-transfer vehicle is configured to releasethe payload satellite when the orbit-transfer vehicle is maneuvered tothe target orbit.